Power safety instrument system

ABSTRACT

A power safety system is configured to provide power information in an aircraft. The power safety system includes a power safety instrument having a power required indicator and a power available indicator, each being located on a display. A position of the power required indicator and the power available indicator represent the power available and power required to perform a hover flight maneuver. The power safety system may be operated in a flight planning mode or in a current flight mode. The power safety system uses at least one sensor to measure variables having an effect on the power required and the power available.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. application Ser. No.13/641,325, filed 30 Jan. 2013, titled “Power Safety Instrument System,”which is a National Stage Entry of P.C.T. Application No.PCT/US2010/061701 filed 22 Dec. 2010, titled “Power Safety InstrumentSystem,” both of which are hereby incorporated by reference for allpurposes as if fully set forth herein. Also, U.S. Pat. No. 7,414,544 toOltheten et al., is hereby incorporated by reference.

TECHNICAL FIELD

The system of the present application relates to a flight instrument foran aircraft. In particular, the system of the present applicationrelates to a power safety instrument for a rotorcraft.

BACKGROUND

The power required to operate a rotorcraft may substantially changeduring the flight path of the rotorcraft. A rotorcraft typicallyrequires substantially more power during a hover, compared to when therotorcraft is traveling forward at a moderate airspeed. For example, asthe rotorcraft is slowed to a landing, the increased power requirementat hover can consume all the power that the engine(s) have available(particularly with a heavy aircraft at a hot temperature and highaltitude environment) causing loss of rotor rpm, an uncontrolleddescent, and possibly a crash landing. Furthermore, the exact powerrequired during hover is affected by a variety of variables, such aspressure altitude and air temperature.

Typically, a pilot will make pre-flight calculations to predict if therotorcraft will have adequate power available to make an approach tohover. The pilot will typically make these pre-flight calculations bycollecting information from several sources. The calculations mayinclude an estimate of the power required by the aircraft to fly athover at a specific location. Another calculation may include anestimate of the power available by the aircraft at hover at the specificlocation. The power available and power required calculations are thencompared to in order to predict sufficient power margin.

Typically, the aforementioned power available and power requiredcalculations are performed by the pilot on the ground in consultationwith relevant performance charts in the rotorcraft flight manual. If theexpected flight involves performing a hover landing at a landing sitethat is different than the departure site, then the pilot must makeeducated guesses regarding certain conditions at the time of landing.For example, the pilot will typically make an educated guess inpredicting the approximate weight of the rotorcraft at the time oflanding. In addition, the pilot will make an educated guess regardingthe predicted air temperature and pressure altitude at the landing site.Each of these variables can be difficult to accurately predict.

There are many potential error opportunities due to the pilot having toread one or more charts, as well as make predictions regarding futureflight conditions and atmospheric conditions. Furthermore, pilotstypically are very conservative in order to allocate margin for anycalculation errors. As a result, many rotorcraft are not fully utilizedas pilots protect themselves and passengers from small but consequentialerrors and in-flight changes in the predicted variables (aircraftweight, outside air temperature, and pressure altitude). For example, arotorcraft may make two separate flights transporting passengers from adeparture site to a destination site when in fact the rotorcraft wasfully capable of performing the task in a single flight. Suchunderutilization of rotorcraft cost rotorcraft operators an enormousamount of time and money over the life of a rotorcraft.

Although the developments in rotorcraft flight instrumentation haveproduced significant improvements, considerable shortcomings remain.

BRIEF DESCRIPTION OF THE DRAWINGS

The novel features believed characteristic of the system of the presentapplication are set forth in the appended claims. However, the systemitself, as well as a preferred mode of use, and further objectives andadvantages thereof, will best be understood by reference to thefollowing detailed description when read in conjunction with theaccompanying drawings, in which the leftmost significant digit(s) in thereference numerals denote(s) the first figure in which the respectivereference numerals appear, wherein:

FIG. 1 is a side view of a rotorcraft according the preferred embodimentof the present application;

FIG. 2 is a partial perspective view of a cockpit portion of therotorcraft from FIG. 1;

FIG. 3 is a schematic view of a power safety system according to thepreferred embodiment of the present application;

FIG. 4 is a schematic view of a calculation of power requiredcalculation, according to the preferred embodiment of the presentapplication; and

FIG. 5 is a schematic view of a calculation of power availablecalculation, according to the preferred embodiment of the presentapplication.

While the system of the present application is susceptible to variousmodifications and alternative forms, specific embodiments thereof havebeen shown by way of example in the drawings and are herein described indetail. It should be understood, however, that the description herein ofspecific embodiments is not intended to limit the method to theparticular forms disclosed, but on the contrary, the intention is tocover all modifications, equivalents, and alternatives falling withinthe spirit and scope of the application as defined by the appendedclaims.

DESCRIPTION OF THE PREFERRED EMBODIMENT

Illustrative embodiments of the system of the present application aredescribed below. In the interest of clarity, not all features of anactual implementation are described in this specification. It will ofcourse be appreciated that in the development of any such actualembodiment, numerous implementation-specific decisions must be made toachieve the developer's specific goals, such as compliance withsystem-related and business-related constraints, which will vary fromone implementation to another. Moreover, it will be appreciated thatsuch a development effort might be complex and time-consuming but wouldnevertheless be a routine undertaking for those of ordinary skill in theart having the benefit of this disclosure.

In the specification, reference may be made to the spatial relationshipsbetween various components and to the spatial orientation of variousaspects of components as the devices are depicted in the attacheddrawings. However, as will be recognized by those skilled in the artafter a complete reading of the present application, the devices,members, apparatuses, etc. described herein may be positioned in anydesired orientation. Thus, the use of terms such as “above,” “below,”“upper,” “lower,” or other like terms to describe a spatial relationshipbetween various components or to describe the spatial orientation ofaspects of such components should be understood to describe a relativerelationship between the components or a spatial orientation of aspectsof such components, respectively, as the device described herein may beoriented in any desired direction.

Referring to FIGS. 1 and 2, a rotorcraft 101 having a power safetysystem 301 is illustrated. Rotorcraft 101 has a fuselage 105 and rotorsystem 103. Rotor system 103 includes a plurality of rotor blades drivenby a rotor mast and at least one engine. In the preferred embodiment,the engine(s) is a turbine engine; however, other engines may be used.Rotor system 103 is configured to provide propulsive forces for flyingin at least a hover mode and a forward flight mode. Rotorcraft 101includes a plurality sensors 309 configured to conduct and record avariety of measurements, such as a fuel gauge, cargo hook load cell, airtemperature gauge, altimeter, engine torque gauge, gas producer speedgauge, gas temperature gauge, engine bleed air indicator, main rotormast torque sensor, fuel flow gauge, and generator load gauge, to name afew. Rotorcraft 101 also includes a cockpit 107 for housing a powersafety instrument 303 of power safety system 301. It should beappreciated that cockpit 107 may any of a variety of cockpit designs,including a “glass cockpit” design in which one or more instruments(such as power safety instrument 303) are digitized and graphicallydisplayed on a screen. It should be appreciated that even thoughrotorcraft 101 is depicted as a helicopter; the scope of the presentapplication is not so limited. For example, rotorcraft 101 may be anyaircraft capable of performing a vertical take-off, a vertical landing,or hover. As such, rotorcraft 101 may be a helicopter, a tiltrotor, atilt-wing aircraft, a hybrid aircraft, or a vertically landing jetaircraft, to name a few. Rotorcraft 101 may also be an unmannedaircraft. An embodiment of power safety system 301 designed for anunmanned aircraft is preferably configured to provide the power safetyinstrument 303 information to a remote operator of the aircraft throughtelemetry, or the like. If the unmanned aircraft does not require aremote operator, then the power safety instrument 303 information may beprocessed directly by the aircraft system responsible for flying theunmanned aircraft.

Referring to FIG. 3, an embodiment of a power safety system 301 isillustrated. Power safety system 301 preferably includes a power safetyinstrument 303, a calculation unit 307, at least one sensor 309, and atleast one data input device 319. Power safety instrument 303 isconfigured for installation in cockpit 107 of rotorcraft 101. Powersafety instrument 303 includes a power gauge 315 visibly portrayed on adisplay 305. Power gauge 315 preferably includes a power required (PR)indicator 311, a power available (PA) indicator 313, and a power usageneedle 317. Power gauge 315 includes numbers 1 through 14, along withrespective hatch marks, which are non-dimensional and are provided forreference only.

Power safety system 301 includes a plurality of sensors, such as sensors309 on rotorcraft 101, which are schematically represented as sensors309 in FIG. 3. Sensors 309 are configured to sense various parameters. Adata input device 319 is configured for manual entry of data. Data inputdevice 319 may be a variety of hardware devices, such as a keyboard, anumeric keypad, twist knob, or a touch screen portion on display 305, toname a few examples. A calculation unit 307 is configured to processdata provided by sensors 309 and data input device 319, as discussedfurther herein. Calculation unit 307 may be any processor capable ofstoring and processing a data for communicating to the pilot via powersafety instrument 303.

Power safety instrument 303 conveniently displays power required (PR)via PR indicator 311, power available (PA) via PA indicator 313, andpower usage via power usage needle 317, all instantaneously calculatedand updated. Thus, the pilot of the rotorcraft is oriented as to howmuch power margin will be available as he brings the rotorcraft into ahover landing. For purposes of this disclosure, power margin is thedifference between PA and PR to operate at a hover, as visually depictedby PA indicator 313 and PR indicator 311, respectively.

It should be appreciated that power safety instrument 303 may take on awide variety of configurations. Outputs from power safety instrument 303may be communicated to the pilot in a variety of methods, includingvisually, audible, and/or through a sensory touch system such as avibration cue, to name a few. Furthermore, even though PR indicator 311and PA indicator 313 are depicted as triangular shapes, each indicator311 and 313 may be configured in a wide variety of shapes and colors.Similarly, power safety instrument 303 may take on a wide variety ofconfigurations. Features of power safety instrument 303 may beintegrated into other instruments within cockpit 107.

Referring now also to FIG. 4, calculation unit 307 calculatesinstantaneous power required (PR) in a PR calculation 401. PR isvisually depicted to the pilot with PR indicator 311 on power gauge 315.PR calculation 401 includes a step 403 for calculating the take-offweight of rotorcraft 101. In one embodiment, the pilot simply calculatesthe take-off weight of rotorcraft 101 by adding the fuel weight,passenger/cargo weight, and any other weight contributing articles. Thepilot then inputs the rotorcraft take-off weight via data input device319. In another embodiment, a landing gear sensor automaticallycalculates the take-off weight and sends the data to calculation unit307. As an added safety measure, PR calculation 401 is configured sothat if the pilot does not enter a rotorcraft weight at take-off, thenthe rotorcraft's maximum gross weight is used as the default. Powersafety system 301 is configured to communicate a message on display 305reminding the pilot to enter and/or verify rotorcraft weight prior totake-off. For example, if a pilot briefly lands rotorcraft 101 totake-on two additional passengers without shutting down, then aweight-on-gear-sensor informs power safety system 301 that therotorcraft has landed, thereby instigating a message on display 305 forthe pilot to enter the new rotorcraft weight prior to take-off.

A step 405 includes accounting for any change of in the weight of fuelin rotorcraft 101. For example, as the rotorcraft burns fuel duringoperation, step 405 includes accounting for changing the currentrotorcraft weight in accordance with the amount of fuel consumed. Theweight of the burned fuel may be calculated by various means, such as afuel flow measuring device, or by simply accounting for changes in thefuel tank gauge. A step 407 includes accounting for changes in a cargohook load. Some rotorcraft may include a cargo hook for supportingexternal loads. As such, a load cell device may be used to measurechanges in the cargo hook load. Other sensors and devices may be used toaccount for changes to the rotorcraft weight during operation. Forexample, if the rotorcraft is a military aircraft with munitions(bullets, missiles, rockets, and the like), the deployment of a munitioncauses the rotorcraft weight to change. A step 406 includes accountingfor any other changes to the weight of the rotorcraft. Furthermore, step406 includes the pilot manually inputting any known change to the weightof the rotorcraft. For example, if the pilot knows he has just lost 500pounds during flight (paratroopers, cargo drop, etc.), then step 406includes the pilot accounting for the weight change with an entry withdata input device 319. Step 406 also includes automatically accountingfor any changes to the rotorcraft weight when such a weight changingactivity is measured or accounted for by any sensor 309.

A step 409 represents the real time weight of the rotorcraft duringoperation, after accounting for weight changes during operation ofrotorcraft 101. As the weight of the rotorcraft decreases, the powerrequired to operate the rotorcraft at hover also decreases. If theweight of the rotorcraft increases during operation (via cargo hook forexample), then the power required to operate the rotorcraft at hoveralso increases.

A step 411 includes continuously measuring outside air temperature (OAT)with an OAT probe, or other temperature measuring device. OAT affectsthe power required to operate the rotorcraft at a hover. For example,generally air at a higher OAT is less dense than air at a lower OAT, fora given altitude. As such, the PR to operate the rotorcraft at hover isa function of OAT. For example, OAT may influence PR due to main rotorblade tip Mach effects. The effect of OAT on PR is preferably determinedusing a look-up table in calculation unit 307. In such an embodiment,rotorcraft specific performance data correlating OAT to PR is stored ina look-up table data format within calculation unit 307, or other datastorage device operably associated with calculation unit 307. In analternative embodiment, OAT measured data may be used to calculate PRusing rotorcraft performance equations.

A step 413 includes continuously calculating the pressure altitude (Hp),which is the air pressure at a particular altitude. Hp affects the powerrequired to operate the rotorcraft at a hover in part because air at ahigher pressure is more compressed and denser, than air at a lowerpressure. As such, the PR to operate the rotorcraft at hover is afunction of Hp. The effect of Hp on PR is preferably determined using alook-up table in calculation unit 307. In such an embodiment, rotorcraftspecific performance data correlating Hp to PR is stored in a look-uptable data format within calculation unit 307, or other data storagedevice operably associated with calculation unit 307. In an alternativeembodiment, Hp measured data may be used to calculate PR usingrotorcraft performance equations.

It should be appreciated that other data may be measured and used in thecalculation of PR. For example, air humidity and wind speed may also beused in the calculation of PR. In certain situations, wind speed may bemeasured at a hover site. The measured wind speed data can becommunicated to the pilot for manual entry via data input device 319.For example, instrumentation at a landing site may measure wind speedand communicate that to the pilot via VHF radio communication.Regardless as to how the pilot is informed of wind speed at the landingsite, the pilot may enter the wind speed data via data input device 319.Additionally, a real-time outside air velocity, measured relative to therotorcraft, may be determined by a low air velocity sensor on therotorcraft, the outside air velocity data being automatically sent tocalculation unit 307. It should be appreciated that alternatively thepilot may input the outside air velocity data into calculation unit 307.The effect of outside air velocity on PR is calculated in calculationunit 307. System 301 may be configured such that outside air velocitydata is used in the calculation of PR only if the outside air velocityexceeds a minimum threshold. For example, a minimum threshold of a 10knot outside air velocity may be used.

A step 421 includes processing the data recorded in steps 409, 411, and413 using rotorcraft hover performance data to derive real-time PR in astep 423. Hover performance data is stored in calculation unit 307. PRmay be derived for at least an in-ground effect (IGE) hover orout-of-ground effect (OGE) hover. OGE can be characterized as performinga hover of the rotorcraft above a distance of approximately one rotordiameter from the ground, or other hard surface. For example, if therotor diameter is 37 feet, then a hover within 37 feet of the groundwould be considered an OGE hover. IGE hover includes a rotorcraft hoverperformed at a certain distance to the ground. The IGE distance istypically defined by a rotorcraft manufacturer as a hover where thelanding gear is a within a certain distance to the ground. IGE hoverrequires less power than OGE due to ground effect influences associatedwith the downwash of the rotor blades causing a high pressure areabetween the rotor blades and the ground. In the preferred embodiment,power safety instrument 303 includes a toggle or other input device sothat the pilot may dictate that the PR be calculated based upon IGE orOGE. Alternatively, IGE or OGE may be automatically determined by one ormore sensors 309 on the rotorcraft 101. It should be appreciated thatcertain landing site conditions may negate ground effect influences uponhover, such as hovering above long grass or water. As such, system 301is configured so that the pilot can dictate PR to be calculated basedupon IGE or OGE.

Furthermore, power safety instrument 303 may be configured tographically communicate to the pilot that PR indicator 311 is beingcalculated based on IGE hover or OGE hover. For example, PR indicator311 may have a “I” character associated with PR indicator 311 tocommunicate to the pilot that PR is currently being calculated basedupon a IGE hover. Similarly, PR indicator 311 may have an “O” characterassociated with PR indicator 311 to communicate to the pilot that PR iscurrently being calculated based upon an OGE hover. As an added safetymeasure, power safety system 301 may be configured to default to OGEsince it requires more power to perform an OGE hover than an IGE hover.As such, if the rotorcraft has enough power to perform an OGE hover,then it has enough power to perform an IGE hover. An example of an IGEhover is taking off or landing at a helipad. An example of an OGE hoveris a logging rotorcraft hovering above a tree line waiting for groundpersonnel to connect a sling load of lumber, the hovering distance beingsuch that PR is calculated using OGE.

Rotorcraft hover performance data is typically supplied by therotorcraft manufacturer in a graphic chart format. Power safety system301 preferably includes the mathematical relationships of the rotorcrafthover performance data in electronic format so that a computer processorin calculation unit 307 may calculate PR based upon the real time datacollection in steps 409, 411, and 413. Step 423 represents the real timecalculation of PR. The real time calculation of PR is graphicaldisplayed on power safety instrument 303 via PR indicator 311.

A step 415 includes a weight-altitude-temperature (WAT) calculation. TheWAT calculation is compared to a WAT limit associated with therotorcraft. The WAT limit of the rotorcraft represents an aircraftlimitation at a certain combination of aircraft weight, altitude, andOAT. Step 415 includes calculating the WAT calculation by analyzing thereal time data acquired in steps 409, 411, and 413. If the WATcalculation is below the rotorcraft WAT limit, then the WAT calculationdoes not produce a limitation, as depicted in step 419. However, if theWAT calculation is equal to or above the rotorcraft WAT limit, then theWAT calculation does produce a limitation, as depicted in step 417. Assuch, when the WAT calculation exceeds the rotorcraft WAT limit, thenpower safety system 301 is configured to communicate that to the pilot.In one embodiment, PR indicator 311 changes size, shape, or color sothat the pilot knows that if he or she attempts to perform a hover atthe current weight/altitude/OAT, then the WAT limit will be exceeded,possibly causing loss of aircraft control during the attempt to performthe hover. The aforementioned transformation of PR indicator 311prevents the pilot from incorrectly thinking it is safe, to perform ahover, take-off, or landing, even though power safety instrument 303 maydepict sufficient margin between PA indicator 313 and PR indicator 311.

Referring now to FIG. 5, calculation unit 307 calculates instantaneouspower available (PA) in a PA calculation 501. PA is visually depicted tothe pilot with PA indicator 313 on power gauge 315. PA calculation 501includes a step 513 for conducting a power assurance check of theengine(s) of rotorcraft 101. Conducting a power assurance check in step513 includes acquiring data from one or more sensors 309 in order toevaluate the health and performance of the engine. A power assurancecheck may be conducted to verify the engine is able to meet minimumrequirements. Furthermore, the power assurance check may be conducted toquantity performance of the engine. Typically, rotorcraft engine(s) areinitially delivered above “min-spec”, meaning the performance of theengine provides more power available than is stated in the rotorcraftspecification and power available charts. For example, if an engine is6% above min-spec, then the engine has approximately 6% more poweravailable than a min-spec engine, at the same ambient conditions.However, unless the pilot is able to account for the 6% above min-specpower, the pilot isn't able to confidently take advantage of the abovemin-spec power available during flight operations. It should also beappreciated that engine performance may degrade over time, thus it isimportant for the pilot to be able to account for engine degradation inorder to continue safe operation of the rotorcraft.

A variety of measurements may be taken in order to conduct the powerassurance check in step 513. For example, a step 503 includes measuringthe torque produced by the engine in the rotorcraft. A step 505 includesmeasuring the OAT which affects the power produced by the engine. A step507 includes measuring the gas producer speed (Ng) of the engine. A step509 includes measuring the pressure altitude (Hp), which affects thepower produced by the engine. A step 511 includes measuring the gastemperature (MGT), which affects the power produced by the engine. Itshould be appreciated that any combination of measurements may be takenin order to conduct the power assurance check in step 513, includingmeasurements other than depicted in steps 503, 505, 507, 509, and 511.

The power assurance check of step 513 is performed periodically, such asin a hover or forward flight procedure. The power assurance check ofstep 513 may also be performed during a pre-flight procedure.Additionally, the power assurance check of step 513 may be conducted bythe pilot, or it may occur autonomously without requiring pilotinteraction. Preferably, calculation unit 307 records a rolling averageof the most recent power assurance checks. For example, the result ofthe ten most recent power assurance checks may be averaged to derive apower assurance calculation. In another embodiment, the power assurancecheck of step 513 may simply provide a pass/fail result, a passingresult meaning that the engine meets the requirements of a min-specengine.

A step 515 includes interpreting the power assurance data from step 513in order to determine the performance of the engine in terms of apercentage above or below min-spec. In one embodiment, power assurancedata from step 513 is compared with min-spec engine power, thus arrivingat percentage above or below min-spec power. In another embodiment, thepower assurance check from step 513 only provides a pass/faildetermination, such that step 515 includes determining a percentageabove or below min-spec engine based upon an input from the pilot. Forexample, if the pilot knows the rotorcraft has a +10% engine (10% abovemin-spec), then the pilot enters that into power safety system 301 viadata input device 319. Alternatively, step 515 may be autonomouslyperformed by calculation unit 307, or manually entered by the pilot.

Preferably, step 515 also includes accounting for any rotorcraftpropulsion configuration which may affect engine performance. Forexample, an air inlet configuration on the rotorcraft may increase ordecrease power available. Furthermore, an air inlet configuration mayinclude a standard inlet, a particle separator inlet, inlet barrierfilter, and snow baffles, to name a few. As such, a particular inletconfiguration may correlate to particular performance data to be used inthe calculation of PA in step 525. Therefore, step 515 includes theability for the pilot to enter information, with data input device 319,regarding a configuration (such as air inlet configuration) that in turnmay dictate particular performance data used in the calculation of PA bycalculation unit 307.

A step 525 includes calculating instantaneous PA based upon theabove/below min-spec engine determination from step 515, as well as aplurality of instantaneous measurements from sensors 309. A step 517includes measuring engine bleed air being drawn from the engine. Themeasurement of engine bleed air usage may include simply determining ifthe engine bleed air switch is on or off. Alternatively, the measurementof engine bleed air may include taking one or more measurements toquantify an amount of bleed air being drawn from the engine. Enginebleed air acts as a drain on engine PA, as such, if engine bleed air isturned off, then PA increases. A step 519 includes measuring OAT, whichhas an effect on PA. A step 521 includes measuring the generator load onthe engine, which affects PA. For example, if the pilot turns off asystem that requires electrical power, then the generator load on theengine decreases, thereby causing the PA to increase. A step 523includes measuring the Hp, which similar to OAT, has an effect on PA. Itshould be appreciated that any combination of measurements may be takenin order to conduct the PA calculation in step 525, includingmeasurements other than depicted in steps 517, 519, 521, and 523. Theinstantaneous calculation of PA in step 525 is graphical displayed onpower safety instrument 303 via PA indicator 313.

In the preferred embodiment, the PA calculated in step 525 is limited byany rotorcraft limitation. For example, in certain cases PA is limitedby a transmission torque limit. In such a situation, PA indicator 313preferably does not exceed the transmission torque limit. Therefore, thepilot can trust that the position of PA indicator 313 is the absolutePA. In one embodiment, PA indicator 313 includes a graphic tocommunicate to the pilot that the PA indicator 313 position is beinglimited by the transmission torque limit, or other rotorcraftlimitation. For example, if the PA indicator 313 is being limited by thetransmission torque limit, then “T” is displayed in the PA indicator313. In alternative embodiment, PA indicator 313 is not restricted bythe transmission torque limit, or other rotorcraft limitation.

In the case of a multi-engine rotorcraft, step 525 may also beconfigured to process data in order to calculate PA for an emergencylanding maneuver upon the loss of an engine. For example, if amulti-engine rotorcraft loses an engine, then PA indicator 313 may beconfigured to represent the PA for the rotorcraft to perform anemergency landing maneuver in a one-engine-inoperable (OEI) condition.As such, the pilot is able to look at power safety instrument 303 anddetermine if there is sufficient margin between the PR indicator 311 andPA indicator 313 in order to perform an emergency landing maneuver atthe current atmosphere conditions and rotorcraft configuration with OEI.Further regarding an OEI condition, if rotorcraft 101 were to lose anengine, then the remaining engine(s) may be required to operate at anunsustainable level. As such, power gauge 315 may be configured so thatPA indicator 313 visually forecast the PA at certain engine levels. Forexample, if an engine is only able to operate 30 seconds at an extremelyhigh level, then the PA indicator 313 may visually communicate that thePA indicator position will decrease after the 30 second time periodexpires. The operable engine may then be able to run at a decreasedlevel for two minutes after the expiration of the 30 second time limit.Therefore, the PA indicator 313 would move accordingly, then visuallycommunicate that the current PA indicator position is valid for the nexttwo minutes. This process may continue if the engine is operating at aPA that is sustainable. Such a feature of system 301 allows the pilot toquickly ascertain margin between PR and PA during an OEI condition foran multi-engine rotorcraft.

Referring again to FIG. 3, power safety system 301 provides forefficient and safe operation of rotorcraft 101. During operation ofrotorcraft 101, power safety system 301 orients the pilot in real-timeso that the pilot can quickly and accurately ascertain whether a hoverlanding or take-off is achievable. Moreover, power safety system 301continuously displays PR and PA at the current rotorcraft location sothat the pilot, by observing the positions of PR indicator 311 and PAindicator 313, may ascertain whether it is safe to perform a hover atany time during the flight. Power safety system 301 also provides theability of the pilot to make changes to the weight of the aircraft inorder to produce sufficient margin between PR and PA to achieve thehover. A non-limiting example is the pilot burning fuel for an extra 30minutes for the sole reason of reducing aircraft weight. Because PR andPA are calculated in real-time, the pilot is able to burn just enoughsurplus fuel as required to result in sufficient margin between PR andPA for the desired hover.

The following non-limiting simplified flight scenario is provided toillustrate power safety system 301 in operation. A rotorcraft pilot istasked with performing a search-and-rescue (SAR) mission for a mountainclimber. The pilot only has a general idea of the area where themountain climber may be located. Because the pilot does not know thefuture hover location, the pilot is unable to make an accuratepre-flight calculation regarding whether the rotorcraft is able toperform a hover at the unknown location. However, the rotorcraft isequipped with power safety system 301. After, the pilot locates themountain climber and a landing area, the pilot simply flies over thehover site while maintaining a sufficient forward speed. Power safetysystem 301 records and processes data necessary to calculate PA and PRto perform a hover at the site. The pilot may also enter the mountainclimber's weight into power safety system 301 via data input device 319.The pilot looks at power safety instrument 303 and ascertains that thereis sufficient margin between PR indicator 311 and PA indicator 313 toperform the hover. The pilot then slows the rotorcraft to a hover andsafely performs the hover at the hover site.

Power safety system 301 is preferably configured to operate in a flightplanning mode, in addition to real-time flight mode. Flight planningmode allows the pilot to predict PA and PR at a future hover location.Power safety system 301 may be operated in flight planning mode eitherpre-flight or during a flight. When power safety system 301 is beingoperated in a flight planning mode during a flight, power safetyinstrument 303 preferably communicates to the pilot that power safetysystem 301 is being operated in a flight planning mode so that the pilotdoes not mistake the positions of PR indicator 311 and PA indicator 313as being real-time flight positions. Flight planning mode operation ofpower safety system 301 involves the pilot manually entering a predictedrotorcraft weight, pressure altitude (Hp), and outside air temperature(OAT), at the desired hover location. It should be appreciated that OATand Hp may be substituted with a density altitude measurement.Furthermore, the pilot may be made aware of hover site conditionsthrough wireless communication, such as VHF radio, from instrumentationat a landing site. Hp and OAT at the desired hover site may beautomatically communicated to power safety system 301 via wireless datatransfer, such as telemetry or the like. For example, if the desiredhover site is an improved landing site with weather instrumentationconnected to transmitter, then power safety system 301 may automaticallyacquire the Hp and OAT data such that the pilot only has to enter apredicted aircraft weight. Operation of power safety system 301 inflight planning mode prevents pilot errors associated with trying toread small paper charts. Furthermore, operation of power safety system301 in flight planning mode allows the rotorcraft to be more fullyutilized by performing the calculations of PA and PR to a high level ofaccuracy.

The flight planning mode is also configured for determining a maximumallowable weight of the rotorcraft while having sufficient marginbetween PA and PR at hover. In such a situation, the pilot enters adesirable amount of margin between PA and PR, and enters the necessarydata. System 301 calculates a maximum weight of the rotorcraft anddisplays that amount to the pilot on display 305.

It should be appreciated that margin between the PA and PR may becommunicated to the pilot in terms of weight margin. For example,display 305 may include a graphic of margin between PA and PR quantifiedin terms of weight margin, the weight margin being the extra amount ofweight the rotorcraft could be carrying while still having sufficient PAto meet the PR to hover. Certain pilots may desire to comprehend marginbetween PA and PR in terms of weight, instead of visually observing adistance between PR indicator 311 and PA indicator 313.

The system of the present application provides significant advantages,including: (1) providing a pilot real-time data during rotorcraftoperation for deciding whether it is safe to perform a hover; (2)allowing a pilot to maximize rotorcraft payload; (3) allowing a pilot tooperate a rotorcraft without having to perform pre-flight hovercalculations; (4) allowing a pilot to determine a hover site whileoperating the rotorcraft; (5) providing a flight planning tool thataccurately calculates predicted power available and power required for apredicted hover; (6) providing a tool that safely determines an amountof weight a rotorcraft can carry while still having sufficient marginbetween PR and PA at the hover site, (7) reducing potential for piloterror; and (8) improving efficiency and safety of the rotorcraft.

The particular embodiments disclosed above are illustrative only, as theapplication may be modified and practiced in different but equivalentmanners apparent to those skilled in the art having the benefit of theteachings herein. Furthermore, no limitations are intended to thedetails of construction or design herein shown, other than as describedin the claims below. It is therefore evident that the particularembodiments disclosed above may be altered or modified and all suchvariations are considered within the scope and spirit of theapplication. Accordingly, the protection sought herein is as set forthin the claims below. It is apparent that a system with significantadvantages has been described and illustrated. Although the system ofthe present application is shown in a limited number of forms, it is notlimited to just these forms, but is amenable to various changes andmodifications without departing from the spirit thereof.

What is claimed is:
 1. A power safety instrument configured to providepower information in an aircraft, the aircraft including a powerplant,the power safety instrument comprising: a display; a calculation unitfor calculating a power required to operate the aircraft in a hoverflight mode, the calculation unit being configured to default tocalculating the power required based upon out-of-ground effect (OGE)performance data; a power required indicator located on the display, thepower required indicator being configured to communicate the powerrequired to operate the aircraft in a hover flight mode; a poweravailable indicator located on the display, the power availableindicator being configured to communicate a power available of thepowerplant in the aircraft; and an input device for allowing a user toinstruct the calculation unit to use in-ground effect (IGE) performancedata for calculating the power required to operate the aircraft in thehover flight mode.
 2. The power safety instrument according to claim 1,wherein the calculation unit calculates the power required based on oneor more variables that each at least partially impact the power requiredto operate the aircraft at a hover.
 3. The power safety instrumentaccording to claim 2, wherein at least one of the variables is at leastone of the following: a current outside air temperature, a currentpressure altitude, and a current aircraft weight.
 4. The power safetyinstrument according to claim 3, wherein the current aircraft weight iscalculated based upon a change in the amount of fuel carried in theaircraft.
 5. The power safety instrument according to claim 3, whereinthe current aircraft weight is calculated based upon a change in a loadcarried by a cargo hook on the aircraft.
 6. The power safety instrumentaccording to claim 3, wherein the current aircraft weight is calculatedbased upon a change in a munitions load carried by the aircraft.
 7. Thepower safety instrument according to claim 1, wherein the calculationunit calculates the power required to operate the aircraft in the hoverflight mode based upon processing at least one measured variable in ahover performance equation.
 8. The power safety instrument according toclaim 1, wherein a position of the power required indicator isrestricted by a weight-altitude-temperature (WAT) calculation.
 9. Thepower safety instrument according to claim 8, wherein the WATcalculation is calculated by an aircraft weight measurement, a pressurealtitude measurement, and an outside air temperature measurement. 10.The power safety instrument according to claim 1, wherein thecalculation unit calculates the power available based on one or morevariables that are each at least partially indicative of the poweravailable of the powerplant in the aircraft.
 11. The power safetyinstrument according to claim 10, wherein at least one of the variablesis at least one of the following: a current outside air temperaturemeasurement, a generator load measurement, a pressure altitudemeasurement, and an engine bleed air measurement.
 12. The power safetyinstrument according to claim 1, wherein the calculation unit calculatesthe power available based upon whether an engine in the powerplant israted differently than a min-spec engine.
 13. The power safetyinstrument according to claim 1, wherein the calculation unit calculatesthe power available based upon a power assurance check, the powerassurance check being configured to test a performance of thepowerplant.
 14. The power safety instrument according to claim 13,wherein the power assurance check results in a pass/fail reading. 15.The power safety instrument according to claim 1, wherein thecalculation unit calculates the power available based in part upon aparticular air inlet configuration.
 16. The power safety instrumentaccording to claim 1, further comprising: a power usage needleconfigured to communicate real-time power usage.
 17. The power safetyinstrument according to claim 1, wherein the power safety instrument isconfigured to operate in a flight planning mode such that a position ofthe power required indicator represents a predicted power required toperform the aircraft in a predicted hover maneuver, and a position ofthe power available indicator represents a predicted power availablefrom the powerplant in the aircraft during the predicted hover maneuver.18. The power safety instrument according to claim 1, wherein a marginbetween a first position of the power required indicator and a secondposition of the power available indicator represents surplus power abovethat required to perform the hover flight mode.